Compressor for a gas turbine

ABSTRACT

A compressor for a gas turbine, in particular an aircraft gas turbine, has a rotor hub carrying rotor blades, a stator equipped with stator vanes, a shroud associated to the stator vanes, and an arrangement providing sealing between the shroud and rotor hub to prevent leakage. To achieve almost complete suppression of leakage air and, concurrently, simplification of design and manufacture, the sealing arrangement ( 20 ) between the shroud ( 8 ) and rotor hub ( 3 ) is formed by a discharge arrangement ( 23 ) for leakage air.

This application claims priority to German Patent Application DE 10 2008 052 101.9 filed Oct. 20, 2008, the entirety of which is incorporated by reference herein.

This invention relates to a compressor for a gas turbine, in particular an aircraft gas turbine, with a rotor hub carrying rotor blades, with a stator equipped with stator vanes, with a shroud associated to the stator vanes, and with an arrangement providing sealing between the shroud and rotor hub to prevent leakage.

The compressor, in particular an axial-flow compressor, includes a rotor shaft with one or several compressor stages having a rotor and a stator. The compressor may be provided with a so-called inlet guide vane assembly upstream of the first stage. The rotor equipped with rotor blades is connected to the compressor or the rotor shaft, respectively, and rotates in the casing of the compressor. The stator of each stage is equipped with stator vanes and is not connected to the compressor shaft. It is therefore stationary. Rotating and non-rotating blade rows alternate in the axial direction of the rotor shaft in a compressor.

Suitable attachment must be provided for the stators. For this, two options exist: 1) The stator has a hub gap, with the rotor shaft extending underneath the stator. In this arrangement, which is found, for example, on high-pressure compressors, the stator is connected to the compressor casing only. 2) On the radially inward side of the stator vanes the stator is provided with a so-called shroud, with the shroud being a ring connecting to the stator at the hub. This arrangement involves that, between rotor and stator hub upstream and downstream of the stator, an axial gap is formed which leads to flow leakage. This leakage must be minimized by a sealing arrangement in the form of seals relative to the rotor shaft. This arrangement is found on the sealing system for a gas turbine according to U.S. Pat. No. 6,932,349 B2. The smaller the gap between seal and shroud, the smaller the leakage flow between the forward and the rearward axial gap.

The arrangement with a hub shroud is mechanically complex, heavy and expensive. Also, the replacement of worn seals incurs considerable assembly effort and cost investment.

A strong transverse duct flow, which interacts with the leakage flow through the shroud and is disturbed, occurs on the stationary hub shroud from the pressure side to the suction side of the stator. The transverse duct flow tends to flow to the vane suction side. As a consequence, flow separation on the stator vane and, thus, severe losses occur. The leakage flow through the axial gap, which is situated before and behind the shroud, is driven by a strong pressure gradient between the leading edge and the trailing edge of the stator and, as it interacts with the main flow, can consequently get large and produce severe losses. Furthermore, heating of the flow is encountered within the shroud cavity. In order to reduce these adverse aerodynamic effects, leakage must be minimized. This can be accomplished by either increasing the number of seals or reducing the gap between seal and shroud. In either case, the geometry is complicated in terms of design and manufacture. As a consequence, increased costs are incurred. Since control of the sealing gap is only difficult to achieve, the operational risk for the compressor is increased as well. Summarizing, then, leakage should be completely prevented as it adversely affects the efficiency of the compressor and, consequently, the fuel consumption of the engine.

It is a particular object of the present invention to avoid these disadvantages of the state of the art.

The present invention, in a broad aspect, provides that the sealing arrangement between shroud and rotor hub is formed by a discharge arrangement for leakage air. Provision is thereby made for an almost complete suppression or elimination of leakage air and, concurrently, a simplification of design and manufacture.

Since special seals are dispensed with in this arrangement, no assembly effort is required for the replacement thereof in the case of wear. This provides for reduced costs and risks as well as improved efficiency and lower fuel consumption.

The present invention is implemented in that the state-of-the-art seals are dispensed with and the leakage flow within the shroud is discharged to the outside through an exhaust cavity within the shroud and the stator vane.

Discharge is accomplished in that the static pressure applied to the exhaust cavity is lower than the pressure prevailing in the shroud cavity. The static pressure must here be that much lower than the pressure prevailing in the shroud cavity, that the desired amount of reduced mass flow is removed.

This can be accomplished, for example, in that the exhaust cavity is connected via suitable lines to either an upstream compressor duct or to a secondary air system of the engine, with the removed air being used as turbine cooling air or for the pressurization of the aircraft cabin or other engine or aircraft systems, for example.

The amount of the removed, absolute leakage mass flow is controllable in that a defined throat is provided within the exhaust cavity or within the discharge arrangement.

Complete suppression of the leakage flow between the rearward and the forward axial gap requires that the shroud geometry, more precisely the flow-wetted area of the shroud, be appropriately attuned. The required area can be determined by way of the pressures before and behind the shroud and the exhaust mass flow.

Generally, the leakage mass flow through a shroud cavity with seals amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.

By way of the exhaust cavity, both the boundary layer before the stator and the boundary layer behind the stator are ingested into the shroud cavity and finally discharged to the outside. This leads to a significant reduction of the stator losses and, thus, to an increase in compressor efficiency. With use being also made of the discharged air, further potential for lowering fuel consumption is provided.

The exhaust cavity itself can have any cross-sectional shape and be provided, for example, as discrete round, elliptical or polygonal openings in any number and position on the shroud. Within the shroud, the cavity can either remain discrete or expand into a circumferential cavity. Vane passage can be provided by one or several tubes or by hollowing the vane, for example.

The shroud without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals. This provides for weight and cost savings. Furthermore, the leakage flow, which entails both losses and heating of the fluid and the components, is nonexistent. Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses. Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable.

The present invention is more fully described below, in light of the accompanying figures showing an embodiment of a compressor for a gas turbine, here an aircraft gas turbine:

FIG. 1 shows a schematic longitudinal section through a compressor in accordance with the state of the art,

FIG. 2 is a schematic representation of the flow phenomena on a compressor in accordance with the state of the art as per FIG. 1,

FIG. 3 shows a schematic longitudinal section through a compressor in accordance with the present invention,

FIG. 4 shows the air discharge on a compressor in accordance with the present invention as per FIG. 3,

FIG. 5 is a representation of the air discharge on a compressor in accordance with the present invention,

FIG. 6 shows a variant of the air discharge on a compressor in accordance with the present invention,

FIG. 7 is a schematic representation of the discharge sections of the leakage air on the shroud,

FIG. 8 is a representation, analogically to FIG. 7, with other discharge sections, and

FIGS. 9 a to c show different air passage sections for the leakage air in a stator vane.

FIG. 1 shows, in a schematic longitudinal section through a stage of a compressor according to the state of the art, a rotor 1 with rotor shaft 2, rotor hub 3 and a rotor blade 4 as well as a stator 5 with casing 6 and a stator vane 7 with a shroud 8 encompassing all stator vanes 7 of the compressor stage on the radially inward side. Disposed between the rotor blade 4 and the stator 5 is a sealing gap 9. For accommodating the shroud 8, a radially extending shroud cavity 11 is provided in the rotor hub 3 which clears a forward axial gap 12 and a rearward axial gap 13 in the axial direction of the rotor shaft 2 and towards the rotor hub 3 thereof. In the radial clearance 14 between the radially inner sealing surface 15 a of the shroud 8 and the hub-side bottom 15 b of the clearance 14, two cross-sectionally frustum-shaped seals 16 and 17 are provided as sealing arrangement 10, each forming a radial gap 18 to the sealing surface 15 a of the shroud 8.

In this arrangement, a leakage flow occurs between the rotor hub 3 and the shroud 8 of the stator 5 in the forward and the rearward axial gap 12, 13, which is reduced by the two seals 16, 17 relative to the rotor shaft 2. The smaller the radial gap 18 between the seals 16, 17 and the shroud 8, the lower the flow leakage between the forward and the rearward axial gap 12, 13.

As illustrated in FIG. 2, with the rotor 1 moving in rotary direction 19, a strong transverse duct flow 21 occurs on the stationary shroud 8 of the stator vanes 7 from the pressure side D to the suction side S of the stator 5 which interacts with the leakage flow 22 through the shroud 8 and is disturbed. The transverse duct flow 21 tends to flow to the suction side S of the stator vane 7. This leads to flow separation on the stator vane 7 and consequentially to severe losses. The leakage flow 22 through the axial gap 12, 13, which is situated before and behind the shroud 8, is driven by a strong pressure gradient between the leading and the trailing edge of the stator 5 and, as it interacts with the main flow, can consequently get large and produce severe losses. Furthermore, heating of the flow is encountered within the shroud cavity 11. In order to reduce these adverse aerodynamic effects, leakage must be minimized. This can be accomplished by either increasing the number of seals 16, 17 or reducing the radial gap 18 between seal 16, 17 and shroud 8. In either case, the geometry is complicated in terms of design and manufacture. As a consequence, increased costs are incurred. Since control of the radial gap 18 is difficult to achieve, the operational risk for the compressor is increased as well. Summarizing, then, leakage should be completely prevented as it adversely affects the efficiency of the compressor and, consequently, the fuel consumption of the engine.

FIGS. 3 to 9 show the aircraft gas-turbine compressor according to the present invention with rotor 1, rotor shaft 2, rotor hub 3, rotor blade 4, stator 5, casing 6, stator vane 7, shroud 8, sealing gap 9, shroud cavity 11, forward and rearward axial gap 12, 13 and clearance 14 with hub-side bottom 15 b.

Provided between the shroud 8 and the rotor hub 3 as sealing arrangement 20 is a discharge arrangement 23 for the leakage air which is removed from the shroud cavity 11 via an air discharge opening 24 and an air discharge duct 25 through the shroud 8 and the stator vane 7. The air discharge duct 25 here forms an exhaust cavity 30 for leakage air.

This arrangement provides for almost complete suppression of leakage air and, concurrently, simplification of design and manufacture. Since special seals 16, 17—as in the state of the art according to FIG. 1—are here dispensed with, no assembly effort for a replacement thereof is required in the case of wear. This provides for a reduction of costs and risks as well as improved efficiency and, thus, reduced fuel consumption of the aircraft gas turbine using the compressor according to the present invention.

The embodiment according to the present invention is implemented in that the state-of-the-art seals 16, 17 in accordance with FIG. 1 are dispensed with and the leakage flow 22 from the shroud cavity 11 is discharged to the outside through the exhaust cavity 30 of the discharge arrangement 23 within the shroud 8 and the stator vane 7.

Discharge is accomplished as per FIG. 5 in that the static pressure p_(a) applied to the exhaust cavity 30 in the form of the air discharge duct 25 is lower than the pressure p_(d) prevailing in the shroud cavity 11. The static pressure p_(a) must here be sufficiently lower than the pressure pa prevailing in the shroud cavity 11, that the desired amount of reduced mass flow is removed.

As shown in FIG. 4, this can for example be accomplished in that the exhaust cavity 30 is connected via suitable return lines 27 to either an upstream compressor duct 26 or to a secondary air system 28 of the engine, with the removed air being used as turbine cooling air or for the pressurization of the aircraft cabin or other engine or aircraft systems, for example.

The amount of the removed, absolute leakage mass flow is controllable in that a defined throat 29 (FIG. 4) is provided within the exhaust cavity 30 or within the discharge arrangement 23.

Complete suppression of the leakage flow 22 between the rearward and the forward axial gap 13, 12 requires that the shroud geometry, more precisely the flow-wetted area 40 of the shroud 8, be appropriately attuned. The required area 40 (FIG. 3) can be determined by way of the pressures p₁ and p₂ of the ingested boundary layers 31 or 32, respectively, before and behind the shroud 8 and the exhaust mass flow (FIG. 5).

In accordance with the state of the art, the leakage mass flow through the shroud cavity 11 with seals 16, 17 generally amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity 30 can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.

By way of the exhaust cavity 30, both the boundary layer 31 before the stator 7 and the boundary layer 32 behind the stator 7 are ingested into the shroud cavity 30 and finally discharged to the outside (FIG. 5). This leads to a significant reduction of the stator losses and, thus, to an increase in compressor efficiency. With use being also made of the discharged air, further potential for lowering fuel consumption of the engine, is provided.

A variant of the air discharge is shown in FIG. 6. While in the arrangement according to FIGS. 3 to 5 the air discharge opening 24 is disposed in the radial end face of the shroud 8, FIG. 6 shows two air discharge openings 41, 42 which, arranged in the axial end faces of the shroud 8, form the exhaust cavity 30 connected to the air discharge ducts 25 in the shroud 8.

The exhaust cavity 30 itself can have any cross-sectional shape. See FIGS. 7 and 8, for example, showing discrete polygonal, round or elliptical openings 33, 34 or 35, respectively, in any number and position on the shroud 8. Within the shroud 8, the exhaust cavity 30 can either remain discrete or expand into a circumferential cavity. Passage through the stator vane 7 can be provided by one or several tubes as round hole 36, oblong hole 37 or slot 38, or by hollowing the stator vane 7 in the form of a hollow chamber 39 (FIGS. 9 a-c), for example.

The shroud 8 without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals 16, 17 in accordance with the state of the art as per FIGS. 1 and 2. This provides for weight and cost savings. Furthermore, the leakage flow, which entails both losses and heating of the fluid and the components, is nonexistent. Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses. Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption of the engine. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable.

LIST OF REFERENCE NUMERALS

-   1 Rotor -   2 Rotor shaft -   3 Rotor hub -   4 Rotor blade -   5 Stator -   6 Casing -   7 Stator vane -   8 Shroud -   9 Sealing gap -   10 Sealing arrangement -   11 Shroud cavity -   12 Axial gap -   13 Axial gap -   14 Clearance -   15 a Sealing surface of shroud 8 -   15 b Hub-side bottom of clearance 14 -   16 Seal -   17 Seal -   18 Radial gap -   19 Rotary direction -   20 Sealing arrangement -   21 Transfer duct flow -   22 Leakage flow -   23 Discharge arrangement -   24 Air discharge opening -   25 Air discharge duct -   26 Compressor duct -   27 Return line -   28 Secondary airflow -   29 Throat -   30 Exhaust cavity -   31 Boundary layer -   32 Boundary layer -   33 Polygonal opening -   34 Round opening -   35 Elliptical opening -   36 Round hole -   37 Oblong hole -   38 Slot -   39 Hollow chamber -   40 Flow-wetted area -   41 Air discharge opening -   42 Air discharge opening -   D Pressure side -   S Suction side 

1. A compressor for a gas turbine, comprising: a rotor hub carrying rotor blades; a stator having stator vanes; a shroud associated with the stator vanes; and an arrangement providing sealing between the shroud and rotor hub to prevent leakage, the sealing arrangement being formed by a discharge arrangement for leakage air.
 2. The compressor of claim 1, wherein the discharge arrangement is an exhaust cavity positioned within the shroud and the stator vanes, to discharge the leakage air to the outside.
 3. The compressor of claim 2, wherein a static pressure applied to the exhaust cavity is lower than a pressure prevailing within a shroud cavity.
 4. The compressor of claim 3, further comprising a return line connecting the exhaust cavity to an upstream compressor duct.
 5. The compressor of claim 3, wherein the exhaust cavity is connected to a secondary air system of the engine.
 6. The compressor of claim 5, and further comprising a defined throat positioned within at least one of the exhaust cavity and the air discharge duct of the discharge arrangement to regulate flow.
 7. The compressor of claim 6, wherein a flow-wetted area of an air discharge opening of the air discharge arrangement is determined by pressures of boundary layers ingested before and behind the shroud or in the shroud cavity, respectively.
 8. The compressor of claim 7, wherein at least one of the exhaust cavity and a cross-section of the air discharge opening in the shroud is provided as a discrete polygonal, round or elliptical opening.
 9. The compressor of claim 7, wherein the exhaust cavity in one of the stator vanes is provided as at least one of a round hole, an oblong hole, a slot and a hollow chamber.
 10. The compressor of claim 1, wherein a static pressure applied to the exhaust cavity is lower than a pressure prevailing within a shroud cavity.
 11. The compressor of claim 1, and further comprising a return line connecting the exhaust cavity to an upstream compressor duct.
 12. The compressor of claim 1, wherein the exhaust cavity is connected to a secondary air system of the engine.
 13. The compressor of claim 1, and further comprising a defined throat positioned within at least one of the exhaust cavity and the air discharge duct of the discharge arrangement to regulate flow. 